1. Field of the Invention
The present invention relates generally to air cooled turbine airfoils, and more specifically to the cooling of a turbine airfoil leading edge.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a turbine section includes a plurality of stages of stator vanes and rotor blades to convert chemical energy from a hot gas flow into mechanical energy by driving the rotor shaft. The engine efficiency can be increased by passing a higher gas flow temperature through the turbine section. The maximum temperature passed into the turbine is determined by the first stage stator vanes and rotor blades.
These turbine airfoils (stator vanes and rotor blades) can be designed to withstand extreme temperatures by using high temperature resistant super-alloys. Also, higher temperatures can be used by providing internal convection cooling and external film cooling for the airfoils. Complex internal cooling circuits have been proposed to maximize the airfoil internal cooling while using a minimum amount of pressurized cooling air to also increase the engine efficiency.
Besides allowing for a higher external temperature, cooling of the airfoils reduces hot spots that occur around the airfoil surface and increase the airfoil oxidation and erosion that would result in shorter part life. This is especially critical in an industrial gas turbine engine where operation hot times between engine start-up and shut-down is from 24,000 to 48,000 hours. Unscheduled engine shut-down due to a damaged part such as a turbine airfoil greatly increases the cost of operating the engine.
In a gas turbine engine, especially in an industrial gas turbine engine, the first stage stator vanes and rotor blades are exposed to the highest gas flow temperatures in the turbine section. The leading edge region of these airfoils is exposed directly to the hot gas flow. One prior art showerhead film cooling arrangement is disclosed in U.S. Pat. No. 7,114,923 B2 issued to Liang on Oct. 3, 2006 and entitled COOLING SYSTEM FOR A SHOWERHEAD OF A TURBINE BLADE, where the airfoil leading edge is cooled by a showerhead arrangement of film cooling holes that include three rows extending along the airfoil in the spanwise direction. See FIGS. 1 and 2. The middle film row is positioned at the airfoil stagnation point where the highest heat load is located. Film cooling holes for each film row are at inline pattern and incline at from 20 degrees to 30 degrees relative to the blade leading edge radial surface.
Another prior art showerhead arrangement is disclosed in U.S. Pat. No. 6,164,912 issued to Tabbita et al on Dec. 26, 2000 and entitled HOLLOW AIRFOIL FOR A GAS TURBINE ENGINE in which the airfoil leading edge include two rows of film cooling holes each located on opposite sides of a stagnation point, and where each row of cooling holes curves around the airfoil wall in the curvature of the airfoil wall. Each film cooling hole discharges the film cooling air upward and toward the stagnation point of the leading edge.
Fundamental shortfalls associated with the FIG. 1 prior art showerhead arrangement include: the heat load onto the blade leading edge region is parallel to the film cooling hole array which reduces the cooling effectiveness; a portion of the film cooling holes within each film row is positioned behind each other (see FIG. 3) that reduces the effective frontal convective area and conduction distance for the oncoming heat load; and, a realistic minimum film hole spacing to diameter ratio is approximately at 3.0. Below this 3.0 level “zipper effect” cracking may occur for the film row. This translates to a maximum achievable film coverage for that particular film row of around 33% or 0.33 film effectiveness for each showerhead film row.
Despite the variety of leading edge region cooling configurations described in the prior art, further improvement is always desirable in order to allow the use of higher operating temperatures, less exotic materials, and reduced cooling air flow rates through the airfoils, as well as to minimize manufacturing costs.
An object of the present invention is to provide for a turbine airfoil with an improved showerhead film cooling hole geometry that can be used in the blade cooling design, especially for a high temperature blade application with a high leading edge film effectiveness requirement.
Another object of the present invention is to provide for a turbine airfoil with a showerhead film cooling geometry that eliminates the film over-lapping problem of the above cited prior art showerhead arrangements and therefore yield a uniform film layer for the airfoil leading edge region.